In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine. These components include the turbine vanes and turbine blades of the gas turbine, upon which the hot combustion gases directly impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1800-2100° F. These components are subject to damage by oxidation and corrosive agents.
Many approaches have been used to increase the operating temperature limits and service lives of the turbine blades and vanes to their current levels, while achieving acceptable oxidation and corrosion resistance. The composition and processing of the base materials themselves have been improved. Cooling techniques are used, as for example by providing the component with internal cooling passages through which cooling air is flowed.
In another approach used to protect the hot-section components, a portion of the surfaces of the turbine blades is coated with a protective coating. One type of protective coating includes an aluminum-containing protective coating deposited upon the substrate material to be protected. The exposed surface of the aluminum-containing protective coating oxidizes to produce an aluminum oxide protective layer that protects the underlying substrate.
Different portions of the gas turbine blade require different types and thicknesses of protective coatings, and some portions require that there be no coating thereon. The application of the different types and thicknesses of protective coatings in some regions, and the prevention of coating deposition in other regions, while using the most cost-efficient coating techniques, can pose difficult problems for gas turbine blades which are new-make or are undergoing repair, and may have existing coatings thereon and/or may need new coatings applied. In many cases, it is difficult to achieve the desired combination of protective coatings and bare surfaces. There is a need for an improved approach to such coating processes to achieve the required selectivity in the presence and thickness of the protective coating in some regions, and to ensure its absence in other regions. The present invention fulfills this need, and further provides related advantages.